The present invention generally relates to ceramic-based articles and processes for their production. More particularly, this invention is directed to processes of producing ceramic matrix composite (CMC) components having detailed features, for example, a tip shroud of a turbine airfoil component.
Higher operating temperatures for gas turbines are continuously sought in order to increase their efficiency. Though advances in Nickel-based superalloys have allowed turbines to operate at these higher temperatures, a step change in capability exists in alternative materials currently being investigated. Ceramic materials are a notable example because their high temperature capabilities can significantly reduce cooling air requirements. As used herein, ceramic-based materials encompass homogeneous ceramic materials as well as ceramic matrix composite (CMC) materials. CMC materials generally comprise a ceramic fiber reinforcement material embedded in a ceramic matrix material. The reinforcement material may be discontinuous short fibers dispersed in the matrix material or continuous fibers or fiber bundles oriented within the matrix material. The reinforcement material serves as the load-bearing constituent of the CMC in the event of a matrix crack. In turn, the ceramic matrix protects the reinforcement material, maintains the orientation of its fibers, and serves to dissipate loads to the reinforcement material. Silicon-based composites, such as silicon carbide (SiC) as the matrix and/or reinforcement material, are of particular interest to high-temperature applications, for example, high-temperature components of gas turbines including aircraft gas turbine engines and land-based gas turbine engines used in the power-generating industry. Continuous fiber reinforced ceramic composites (CFCC) are a particular type of CMC that offers light weight, high strength, and high stiffness for a variety of high temperature load-bearing applications, including shrouds, combustor liners, vanes (nozzles), blades (buckets), and other high-temperature components of gas turbines. A notable example of a CFCC material developed by the General Electric Company under the name HiPerComp® contains continuous silicon carbide fibers in a matrix of silicon carbide and elemental silicon or a silicon alloy.
Examples of CMC materials and particularly SiC/Si-SiC (fiber/matrix) CFCC materials and processes are disclosed in U.S. Pat. Nos. 5,015,540, 5,330,854, 5,336,350, 5,628,938, 6,024,898, 6,258,737, 6,403,158, and 6,503,441, and U.S. Patent Application Publication No. 2004/0067316. One such process is known as “prepreg” melt-infiltration (MI), which in general terms entails the fabrication of CMCs using multiple prepreg layers, each in the form of a tape-like structure comprising the desired reinforcement material, a precursor of the CMC matrix material, and one or more binders.
For purposes of discussion, a low pressure turbine (LPT) blade 10 of a gas turbine engine is represented in FIG. 1. The blade 10 is an example of a component that can be produced from a ceramic-based material, including CMC materials. The blade 10 is generally represented as being of a known type and adapted for mounting to a disk or rotor (not shown) within the turbine section of an aircraft gas turbine engine. For this reason, the blade 10 is represented as including a dovetail 12 for anchoring the blade 10 to a turbine disk by interlocking with a complementary dovetail slot formed in the circumference of the disk. As represented in FIG. 1, the interlocking features comprise protrusions referred to as tangs that engage recesses defined by the dovetail slot, though other interlocking features can be used. The blade 10 is further shown as having a platform 14 that separates an airfoil 16 from a shank 18 on which the dovetail 12 is defined. The blade 10 is further equipped with a blade tip shroud 20 which, in combination with tip shrouds of adjacent blades within the same stage, defines a band around the blades that is capable of reducing blade vibrations and improving airflow characteristics. By incorporating a seal tooth 22, the blade tip shroud 20 is further capable of increasing the efficiency of the turbine by reducing combustion gas leakage between the blade 10 and a shroud surrounding the blade tip. The tip shroud 20 has very demanding material requirements because it is directly subjected to hot combustion gases during operation of the engine and high centrifugal loading.
Current state-of-the-art approaches for fabricating ceramic-based turbine blades have involved integrating the dovetail 12, platform 14, airfoil 16 and tip shroud 20 as one piece during the manufacturing process, much like conventional investment casting techniques currently used to make metallic blades. However, the tip shroud 20 (along with the dovetail 12 and platform 14) represents a detailed geometric feature of the blade 10 that poses substantial challenges to designing, manufacturing and integrating CMC components into an affordable, producible design for turbine applications. For example, the process of integrating the tip shroud 20 with the airfoil 16 using CMC materials creates complexities in the design and manufacturing process, and can result in a process that can be too expensive to be economically practical. Furthermore, the low strain-to-failure capabilities of typical CMC materials pose additional challenges to implementing CMC materials in shrouded blade designs.